The present disclosure relates to gas turbine engines, and more particularly to a fuel injector therefor.
Gas turbine engines, such as those which power modern military aircraft, include a compressor section to pressurize a supply of air, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases and generate thrust. Downstream of the turbine section, an augmentor section, or “afterburner”, is operable to selectively increase the thrust. The increase in thrust is produced when fuel is injected into the core exhaust gases downstream of the turbine section and burned with the oxygen contained thereinto generate a second combustion.
Typically, the injected fuel is controlled to penetrate relatively deep into the core exhaust gases to increase augmentor efficiency as well as the magnitude of the supplemental engine thrust. Such deep fuel penetration, however, is dependent on the fuel flow rate which may negatively impact flame stability and increase augmentor instabilities commonly called “screech”.
Traditional fuel flow distribution control in the augmentor section may include additional fuel circuits, additional injection sites, pressure-acting valves, and/or more complicated injectors that have small passages by necessity; all of which may introduce complexity and cost.